Core compartment vent during engine shutdown to reduced bowed rotor start

This application is a division of U.S. application Ser. No. 18/175,918 filed Feb. 28, 2023, the disclosure of which is incorporated herein by reference in its entirety.

The subject matter disclosed herein generally relates to gas turbine engines and, more particularly, to a method and an apparatus for gas turbine engine system bowed rotor start mitigation and wear reduction.

Gas turbine engines are used in numerous applications, one of which is for providing thrust to an aircraft. Gas turbine engines are typically operated while the aircraft is on the ground, such as during taxiing from a gate prior to takeoff and taxiing back to the gate after landing. Gas turbine engines are typically operated at an idle level to warm engine subsystems, operate accessory subsystems, and keep the aircraft in a ready state. In some instances, on-ground operation at idle can be for extended periods of time, particularly at busy airports. The on-ground operation at idle can result in sustained periods of non-flight fuel burn and contributes to engine wear as well as associated operating noise, tire wear, brake wear, and fuel-burn emissions.

When the gas turbine engine of an airplane has been shut off for example, after an airplane has landed at an airport, the engine is hot and due to heat rise, the upper portions of the engine will be hotter than lower portions of the engine. When this occurs thermal expansion may cause deflection of components within the engine, which may result in a “bowed rotor” condition. If a gas turbine engine is in such a bowed rotor condition, it is undesirable to restart the engine. Engine start in this condition results in rub out of abradable and more open clearances, reducing compressor stability and performance.

The uneven nature of nacelle core compartment temperature results in a similar asymmetric temperature of the engine core cases. The core cases at top remaining hot while bottom structure begins to cool. This results in a level of case bow contributing to nonuniform compressor clearances.

In one embodiment, a core section and nacelle assembly of a gas turbine engine includes a compressor located at an engine central longitudinal axis, a core case enclosing the compressor, and a nacelle located radially outboard of the core case and defining a core compartment between the nacelle and the core case. One or more vent openings are located in the nacelle to circulate a cooling airflow through the core compartment, and one or more fans are positioned at the one or more vent openings to urge the cooling airflow through the one or more vent openings to cool the core compartment.

Additionally or alternatively, in this or other embodiments the one or more fans are positioned inside the nacelle at corresponding vent openings of the one or more vent openings.

Additionally or alternatively, in this or other embodiments a battery pack is operably connected to the one or more fans to power the one or more fans.

Additionally or alternatively, in this or other embodiments the one or more fans are operable to generate electrical power to recharge the battery pack.

Additionally or alternatively, in this or other embodiments the one or more vent openings are two vent openings positioned circumferentially 180 degrees apart relative to the engine central longitudinal axis.

Additionally or alternatively, in this or other embodiments the one or more vent openings are positioned at one or more of a core cowl, an inner fan duct or an upper bifurcation of the nacelle.

Additionally or alternatively, in this or other embodiments a first fan of the one or more fans is located at a first vent opening of the one or more vent openings and is configured to urge cooling airflow into the core compartment via the first vent opening, and a second fan of the one or more fans is located at a second vent opening of the one or mor event openings and is configured to urge cooling airflow out of the core compartment via the second vent opening.

Additionally or alternatively, in this or other embodiments a first fan of the one or more fans is positioned at a first vent opening of the one or more vent openings and is configured to urge cooling airflow out of the core compartment via the first vent opening, a second fan of the one or more fans is positioned at a second vent opening of the one or more event openings and is configured to urge cooling airflow out of the core compartment via the second vent opening.

Additionally or alternatively, in this or other embodiments the one or more fans are selectably operable only when operation of the gas turbine engine is stopped.

In another embodiment, a gas turbine engine includes a combustor to combust a mixture of fuel and air, a turbine located at and driven about an engine central longitudinal axis by gaseous products of the combustion, and a compressor driven by rotation of the turbine. A core case encloses the compressor, and a nacelle is located radially outboard of the core case and defines a core compartment between the nacelle and the core case. One or more vent openings are located in the nacelle to circulate a cooling airflow through the core compartment, and one or more fans are positioned at the one or more vent openings to urge the cooling airflow through the one or more vent openings to cool the core compartment.

Additionally or alternatively, in this or other embodiments the one or more fans are positioned inside the nacelle at corresponding vent openings of the one or more vent openings.

Additionally or alternatively, in this or other embodiments a battery pack is operably connected to the one or more fans to power the one or more fans.

Additionally or alternatively, in this or other embodiments the one or more fans are operable to generate electrical power to recharge the battery pack.

Additionally or alternatively, in this or other embodiments the one or more vent openings are two vent openings positioned circumferentially 180 degrees apart relative to the engine central longitudinal axis.

Additionally or alternatively, in this or other embodiments the one or more vent openings are positioned at one or more of a core cowl, an inner fan duct or an upper bifurcation of the nacelle.

Additionally or alternatively, in this or other embodiments a first fan of the one or more fans is positioned at a first vent opening of the one or more vent openings and is configured to urge cooling airflow into the core compartment via the first vent opening, and a second fan of the one or more fans is positioned at a second vent opening of the one or mor event openings and is configured to urge cooling airflow out of the core compartment via the second vent opening.

Additionally or alternatively, in this or other embodiments a first fan of the one or more fans is positioned at a first vent opening of the one or more vent openings and is configured to urge cooling airflow out of the core compartment via the first vent opening, and a second fan of the one or more fans is positioned at a second vent opening of the one or more event openings and is configured to urge cooling airflow out of the core compartment via the second vent opening.

Additionally or alternatively, in this or other embodiments the one or more fans are selectably operable only when operation of the gas turbine engine is stopped.

In yet another embodiment, a method of cooling a core section of a gas turbine engine includes providing one or more vent openings in a nacelle of a gas turbine engine, the nacelle enclosing a core section of the gas turbine engine including a core case enclosing at least a compressor. One or more fans are provided at each vent opening of the one or more vent openings, and the one or more fans are selectably operated to urge an airflow through the one or more vent openings to ventilate a core compartment defined between the nacelle and the core case. The core section is cooled via ventilation of the core compartment.

Additionally or alternatively, in this or other embodiments the one or more fans are operated only when operation of the gas turbine engine is stopped.

A detailed description of one or more embodiments of the disclosed apparatus and method are presented herein by way of exemplification and not limitation with reference to the Figures.

schematically illustrates a gas turbine engine. The gas turbine engineis disclosed herein as a two-spool turbofan that generally incorporates a fan section, a compressor section, a combustor sectionand a turbine section. Alternative engines might include other systems or features. The fan sectiondrives air along a bypass flow path B in a bypass duct, while the compressor sectiondrives air along a core flow path C for compression and communication into the combustor sectionthen expansion through the turbine section. Although depicted as a two-spool turbofan gas turbine engine in the disclosed non-limiting embodiment, it should be understood that the concepts described herein are not limited to use with two-spool turbofans as the teachings may be applied to other types of turbine engines including three-spool architectures.

The exemplary enginegenerally includes a low speed spooland a high speed spoolmounted for rotation about an engine central longitudinal axis A relative to an engine static structurevia several bearing systems. It should be understood that various bearing systemsat various locations may alternatively or additionally be provided, and the location of bearing systemsmay be varied as appropriate to the application.

The low speed spoolgenerally includes an inner shaftthat interconnects a fan, a low pressure compressorand a low pressure turbine. The inner shaftis connected to the fanthrough a speed change mechanism, which in exemplary gas turbine engineis illustrated as a geared architectureto drive the fanat a lower speed than the low speed spool. The high speed spoolincludes an outer shaftthat interconnects a high pressure compressorand high pressure turbine. A combustoris arranged in exemplary gas turbinebetween the high pressure compressorand the high pressure turbine. An engine static structureis arranged generally between the high pressure turbineand the low pressure turbine. The engine static structurefurther supports bearing systemsin the turbine section. The inner shaftand the outer shaftare concentric and rotate via bearing systemsabout the engine central longitudinal axis A which is collinear with their longitudinal axes.

The core airflow is compressed by the low pressure compressorthen the high pressure compressor, mixed and burned with fuel in the combustor, then expanded over the high pressure turbineand low pressure turbine. The turbines,rotationally drive the respective low speed spooland high speed spoolin response to the expansion. It will be appreciated that each of the positions of the fan section, compressor section, combustor section, turbine section, and fan drive gear systemmay be varied. For example, gear systemmay be located aft of combustor sectionor even aft of turbine section, and fan sectionmay be positioned forward or aft of the location of gear system.

The enginein one example is a high-bypass geared aircraft engine. In a further example, the enginebypass ratio is greater than about six (6), with an example embodiment being greater than about ten (10), the geared architectureis an epicyclic gear train, such as a planetary gear system or other gear system, with a gear reduction ratio of greater than about 2.3 and the low pressure turbinehas a pressure ratio that is greater than about five. In one disclosed embodiment, the enginebypass ratio is greater than about ten (10:1), the fan diameter is significantly larger than that of the low pressure compressor, and the low pressure turbinehas a pressure ratio that is greater than about five 5:1. Low pressure turbinepressure ratio is pressure measured prior to inlet of low pressure turbineas related to the pressure at the outlet of the low pressure turbineprior to an exhaust nozzle. The geared architecturemay be an epicycle gear train, such as a planetary gear system or other gear system, with a gear reduction ratio of greater than about 2.3:1. It should be understood, however, that the above parameters are only exemplary of one embodiment of a geared architecture engine and that the present disclosure is applicable to other gas turbine engines including direct drive turbofans.

A significant amount of thrust is provided by the bypass flow B due to the high bypass ratio. The fan sectionof the engineis designed for a particular flight condition—typically cruise at about 0.8 Mach and about 35,000 feet (10,688 meters). The flight condition of 0.8 Mach and 35,000 ft (10,688 meters), with the engine at its best fuel consumption—also known as “bucket cruise Thrust Specific Fuel Consumption (‘TSFC’)”—is the industry standard parameter of lbm of fuel being burned divided by lbf of thrust the engine produces at that minimum point. “Low fan pressure ratio” is the pressure ratio across the fan blade alone, without a Fan Exit Guide Vane (“FEGV”) system. The low fan pressure ratio as disclosed herein according to one non-limiting embodiment is less than about 1.45. “Low corrected fan tip speed” is the actual fan tip speed in ft/sec divided by an industry standard temperature correction of [(Tram °R)/(518.7° R)] 0.5. The “Low corrected fan tip speed” as disclosed herein according to one non-limiting embodiment is less than about 1150 ft/second (350.5 m/sec).

Referring now to, a core portionof the engineis enclosed in a core case. In some embodiments, the core portionincludes the high pressure compressorand the combustor, and in some embodiments also includes the high pressure turbine. A nacelleof the engineis disposed radially outboard of the core caseand encloses at least the core portionof the engine. The nacelleand the core casedefine a core compartmenttherebetween. The nacelleincludes features to cool the core compartmentwhen the engineis in operation, since the core compartmentmay otherwise accumulate heat which would be detrimental to performance of the engine. More particularly, one or more cooling scoopsare disposed in corresponding nacelle openingsin the nacelle. The cooling scoopincludes a scoop intakethat protrudes from an outer nacelle surfaceinto the bypass flow path B of the engine. The scoop intakeincludes an inlet openingthrough which a portion of the bypass airflowfrom the bypass flow path B is flowed into the cooling scoop. The cooling scoopfurther includes a scoop shaftthat extends through the nacellefrom the scoop intakethrough an inner nacelle surfaceof the nacelleto a scoop outlet. During operation of the engine, the scoop intakeand the scoop shaftdefine a cooling pathwayto direct the portion of bypass airflowas a cooling airflowinto the core compartmentto cool the core compartmentand thereby cool the core case. While in the embodiment ofthe cooling scoopsare located at a core cowlportion of the nacelle, on other embodiments the cooling scoopsmay additionally or alternatively be located at other locations of the nacelle. For example, as illustrated in, the cooling scoopsmay be located at an inner fan ductor, as illustrated in, the cooling scoopsmay be located along an upper bifurcation.

When the engineoperation is stopped, the flow along the bypass flow path B is stopped, but it is still desired to cool the core caseto reduce the incidence of bowed rotor conditions when restarting. When engineoperation is stopped, the core casetends to bow due at least in part to non-uniform temperature distribution in the core compartment. In a bowed rotor state, clearances between rotating airfoils of the high pressure compressorand the core caseare non-uniform and results in interaction or rub of the rotating airfoils and the core caseat one or more circumferential locations around the core case. In the case of bow of the core case, these areas of rub are localized and result in areas of excess operating clearance between the high pressure compressorand the core case. This excess operating clearance in turn causes excess leakage, performance loss and stability loss during operation of the engine.

As such, a fanis disposed in the core compartmentat the scoop outletto circulate the cooling airflowthrough the core compartment. In some embodiments, the fanis powered by a power source, such as a rechargeable battery packoperably connected to a fan motorto drive rotation of the fan. Operation of the fanmay be controlled by a controller, which may initiate operation of the fanbased on commands received from, for example, the aircraft cockpit or from a switch outside of the nacelle that may be activated by ground service personnel. In other embodiments, the controllermay initiate operation of the fanautomatically or based on a command from an electronic engine control (EEC) of the enginewhen the controllerdetects that operation of the enginehas stopped. Further, in some embodiments, during normal operation of the engine, the bypass airflow B flows across the fanurging rotation of the fan, which operates the fan motoras a generator to generate electrical energy utilized to charge the battery pack, or alternatively for use by other engineor aircraft systems or components. In some embodiments, the controllermay utilize programming or control circuits to stop this power generation when the battery packis fully charged.

Referring now to, a plurality of cooling scoopsand corresponding fansare arranged to vent core compartment airflowfrom the core compartmentto outside of the nacelle. For example, in one embodiment illustrated in, cooling scoops, for example, two cooling scoops, are placed at or near a vertical horizon lineof the engine, and under engine stop conditions, fanslocated at each cooling scoopare configured to urge the core compartment airflowout of the core compartment.

In other embodiments, such as illustrated in, the cooling scoopsand corresponding fansare located and arranged to urge the cooling airflowinto the core compartmentvia, for example, a first cooling scoop, and vent core compartment airflowfrom the core compartmentvia a second cooling scoop. A first fanlocated at the first cooling scoopis configured to urge cooling airflowinto the core compartmentvia the first cooling scoop, while a second fanlocated at the second cooling scoopis configured to direct core compartment airflowoutward through the second cooling scoopto outside of the nacelle. In one illustrated embodiment, the first cooling scoopis located at or near a vertical bottomof the nacelle, and the second cooling scoopis located at or near a vertical topof the nacelle. Since after engineshutdown, the heat gradient in the core compartmentis such that the air inside the core compartmentis hotter nearer the top of the core compartment, the configuration of the cooling scoops,and the corresponding fansandforces the hottest core compartment airflowfrom the core compartmentvia the second cooling scoop, while urging cooler core compartment airflowupward from nearer the vertical bottom of the core compartmenttoward the vertical top of the core compartment. This has the effect of both cooling the core compartmentand equalizing a temperature distribution in the core compartment. This thereby reduces a thermal gradient between the vertical top and bottom of the core compartmentto reduce the incidence of bowed rotor conditions.

The term “about” is intended to include the degree of error associated with measurement of the particular quantity based upon the equipment available at the time of filing the application. For example, “about” can include a range of ±8% or 5%, or 2% of a given value.

The terminology used herein is for the purpose of describing particular embodiments only and is not intended to be limiting of the present disclosure. As used herein, the singular forms “a”, “an” and “the” are intended to include the plural forms as well, unless the context clearly indicates otherwise. It will be further understood that the terms “comprises” and/or “comprising,” when used in this specification, specify the presence of stated features, integers, steps, operations, elements, and/or components, but do not preclude the presence or addition of one or more other features, integers, steps, operations, element components, and/or groups thereof.

While the present disclosure has been described with reference to an exemplary embodiment or embodiments, it will be understood by those skilled in the art that various changes may be made and equivalents may be substituted for elements thereof without departing from the scope of the present disclosure. In addition, many modifications may be made to adapt a particular situation or material to the teachings of the present disclosure without departing from the essential scope thereof. Therefore, it is intended that the present disclosure not be limited to the particular embodiment disclosed as the best mode contemplated for carrying out this present disclosure, but that the present disclosure will include all embodiments falling within the scope of the claims.